The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in turbine stages which power the compressor and a shaft that typically drives a fan in an aircraft turbofan engine application.
A high pressure turbine (HPT) directly follows the combustor and receives the hottest gases therefrom from which energy is initially extracted. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the gases.
The first stage turbine nozzle includes radially outer and inner bands bounding a row of hollow nozzle vanes. Various cooling circuits are provided in the nozzle for cooling the various parts thereof to limit temperature and ensure long life.
The nozzle is typically mounted in the engine by an integral mounting flange extending inwardly from the inner band.
The outer band of the nozzle is suitably configured to bound the hot combustion gas flow between the outlet end of the combustor and the surrounding turbine shroud of the first stage turbine rotor blades.
The outer band includes a forward radial flange and a corresponding seal joining the outer liner of the combustor. An aft radial flange axially adjoins the hanger supporting the turbine shroud and includes a W-seal in an exemplary configuration.
This aft seal is axially compressed between a seat in the aft flange and the hanger, and substantial axial compression preloads are maintained through the aft flange upon assembly.
The aft flange therefore is relatively stiff to support reaction forces from the shroud hanger, and such stiffness is typically provided by a relatively thick aft flange, particularly at its root junction with the aft end of the outer band.
The thick aft flange increases the thermal mass thereof and increases the difficulty of cooling during operation.
Cooling is typically provided in the turbine nozzle by channeling compressor discharge pressure (CDP) air over the outer band with portions thereof being channeled through internal chambers of the vanes.
The outer band may include arrays of film cooling holes distributed therethrough. The aft flange itself may include cooling holes extending axially aft therethrough.
However, any cooling air directly channeled through the aft flange bypasses the turbine nozzle vanes and decreases nozzle performance as chargeable air.
Conversely, cooling air which is discharged inwardly through the outer band is therefore recovered between the nozzle vanes and retains performance as non-chargeable air.
The complex aft flange of the nozzle outer band typically requires trade-offs in design for carrying the high preloads within acceptable stress limits, while also being adequately cooled for nozzle life.
However, the thick aft flange limits the ability to effectively cool the flowpath, and experience has shown that insufficient cooling leads to thermal distress of the aft flange and outer band thereat which reduces nozzle life.
Accordingly, it is desired to provide a turbine nozzle having an outer band with improved cooling of the aft flange using non-chargeable flow.